Tailless airplane



Apil 6, 1948.V

' w; H. KoRFF TAILLEss AIRPLANE 4 sheets-sheet 1 Fild oct. 5, 1942 MUC-PH. KQRFF April 6, 1948. -w. H. yKom-F TAILLEss AIRPLANEY Filed- 0G13'.5, 1942 4 Sheets-Sheet 2 SEM/SPAN IN VENTO/P WAL TER H. KOR/ff w. H.KORFF TAILLESS AIRPLANE April 6, 1948.

Filed OGL@ 1942 ANGLE OF ATAC/f,CZ,DL`G., 0.20 PLA/N FLA/D FLP L'NGT/'-IN1/[N709 PERCENT Of .YM/SPAN FROM CENTER WALTER H. Kopf-F Patented Apr.6, 1948 TAILLESS AIBPLANE Walter H. Korff, Burbank, Calif., assignor toLockheed Aircraft Corporation, Burbank, Calif.

Application October 5, v1942, Serial No. 460,797

4 Claims. (Cl. 244-13) This invention relates to improvements in theso-called tailless airplane of the type having a single main supportingairfoil and without the conventional horizontal stabilizer airfoilusually employed in a position either to the rear of or in front of themain airfoil.

In the ordinary airplane of conventional design employing a rearwardlypositioned horizon tal stabilizer maintenance of stability and controlrequire that a downward load be carried upon the stabilizer particularlyunder conditions of high angles of attack such as at take-ois andlanding, and this down load in addition to the gross weight of theairplane must necessarily be borne by the lift of the main wing.Consequently, in this type of airplane, for a given load carryingcapacity either the landing or take-off speeds or the main wing area orboth must be increased to provide the added lift equal to the horizontalstabilizer down load and in either case additional power consumption isrequired for a` given pay load. The rearwardly located horizontalstabilizer arrangement also has the disadvantage of increasing the totalstructural weight, the aerodynamic drag and the cost of construction ofthe airplane as a whole.

In the so-called forward tail type of airplane in which the horizontalstabilizer is located forward of the main wing, an upward load isnormally carried on the stabilizer assisting the main wing in supportingthe airplane, particularly at high angle of attack conditions.Consequently, in this forwardly located horizontal stabilizer type ofairplane, lower landing and take-olf speeds for a given load or giventotal wing area are possible and moreover, under all conditions lesspower is required for flight as compared to the conventional airplanehereinbefore mentioned. The forwardly located horizontal stabilizer,however, as in the case of the rearwardly located stabilizer, still hasthe disadvantage of added aerodynamic drag, structural weight andcomplication and cost of construction.

In the so-called tailless type of airplane as heretofore constructed,the disadvantages of drag, weight and cost associated with the forwardor rearward horizontal stabilizer arrangement have been largelyeliminated with the omission of the separate stabilizer surfaces, butstability of the airplane and control under high angle of attackconditions with this arrangement has heretofore been accomplishedlargely by employing sweep-back in the wings and providing means tomaintain a down load on the trailing wing tips which act in the samemanner and serves the same purpose as the down load on the conventionalrearwardly located stabilizer of the conventional airplane ashereinbefore described. Under such conditions the effective coefficientof lift of the wing as a whole is low, particularly at high angles ofattack. For' this and other reasons the tailless type of airplane, todate, has shown few, if any, advantages over aircraft of goodconventional design. However, this tailless design, which virtuallyembodies main wing and rearwardly located stabilizer in one single wingof continuous and increased span, has the advantage over both theforward and rearward tail types of construction in resulting in anoverall reduction in the effective induced and profile drag and weightfor a given total airfoil area.

Objects of the present invention are to overcome and eliminate thehereinbefore mentioned aerodynamic and structural disadvantages inherentin the three types of airplanes hereinbefore described and to takeadvantage of and combine in a single airplane of simplified constructionall of the advantages inherent in both the forward tail and taillesstypes of airplane construction.

The objects of this invention are realizedin an embodiment of thisinvention in which the desirable features of both the forward stabilizertype and so-called tailless type of airplanes are com-- bined and thedisadvantages associated with these two types of airplanes are largelyeliminated, resulting in a tailless airplane having the characteristicsand advantages of a forward stabilizer type of airplane in regard toeiiiciency of use of the entire airfoil surfaces for producing usefullift under all conditions and attitudes of the airplane and having theaerodynamic and structural advantage associated with utilizing a singlesustaining airfoil of increased aspect ratio andthe elimination of aseparate stabilizer.

The invention broadly stated resides in a single longitudinallystableairfoil, the entire surface of which normally and under suchconditions of stability, carries a positive lift.

The invention more specifically stated resides in a singlelongitudinally stable airfoil carrying sweep-back and aerodynamic orgeometrical twist or wash-out in the laterally extendingpor# tionsthereof of such degree and so combined as to result in a forward travelof the center of pressure with decreasing angles of attack and arearward travel of the center of pressure with increasing angles ofattack for all normal flight conditions while at the same timemaintaining positive lift on all portions of such airfoil.

The invention further resides in an airplane employing the beforementioned inherently stable airfoil, combining sweep-back and wash-outand additionally an elevator control flap constituting a hinged portionof the trailing edge adjacent the rootsection thereof, the length andposition of such hinged flap being such that lowering of the flapresults in simultaneous increase of the effective maximum coeflicientofy lift and a forward movement of the center of pressure of the airfoilas a whole resulting in a stalling moment whereby high angles of attacklof the airfoil with simultaneous increase in the overall effectivecoefficient and maximum coefcient of lift may be attained.

rIhe invention in its narrower aspects resides in an airplane employingthe before mentioned' inherently stable airfoil combining sweep-back andWash-out and additionally, elevator control trailing edge flaps adjacentthe root section, ailerons adjacent the trailing airfoil tips and liftmodifying flaps located intermediate the elevator flaps and aileronswhereby high angle of attack, increased maximum coefficient of lift andincreased drag or braking action can be simultaneously effected.

Other objects and features of novelty will be evident hereinafter.

In the drawings in which preferred embodiments .of the invention areillustrated;

Figure 1 is a semi-plan view of the airplane which is symmetrical aboutthe center line, showing the general arrangement of the fuselage wingsand control surfaces;

Figure 2 is a front elevation of the airplane, of Figure 1;

Figure 3 is a side elevation of the airplane of Figure 1;

Figure 4 is a cross-section of the wing taken at 4-4 of Figure 1;

Figure 5 is a cross-section of the wing taken at 4 4 of Figure 1 showingan optional type of flap:

Figure 6 is a plan view of the airplane similar to Figure 1 butillustrating a modified form of the invention;

Figures 7A and 7B are plan diagrams of wings equivalent to that of thepresent invention eX- cept without sweep-back or wash-out illustratingcertain aspects of the derivation of the conditions for stability andcontrol.

Figures 8A and 8B are plan diagrams of the wing illustrating thederivation of the loci of center of pressure travel and the properdegree of sweep-back.

Figure 9 is a graphical representation of the lift and center ofpressure characteristics of a typical airfoil section prole of the typeemployable in the present invention.

Figure 10 is a loading diagram for the semispan of the wing with theelevator flap at several positions.

Figure 11 is a characteristic curve showing the relationship of flapspan length to flap control moment;

Referring now to the drawings throughout which the same referencecharacters refer to similar parts, I0 is a suitable fuselage carrying alaterally extending main wing, the semi-span portion extending to theleft of line II of the longitudinal plane of symmetry which is shown atI2, having a sweep-back angle with respect to the 50% chord center lineI3. Vertical stabilizers and rudders are carried at the wing tips asshown at I4 and I5 respectively and the trailing edge of the said wingcarries a pair of plain hinged or split, partial span flaps as shown atI8, each adapted to be raised or lowered through suitable angles aboutthe hinge line il as best shown at i3 and 29 respectively in Figures 4and 5.

In Figure 6, in which anoptional form of the invention is illustrated,the partial span flaps i3 and ailerons I9 which are adjacent thefuselage and wing tips respectively, are made shorter thanthose shown inFigure 1 to provide for an additional intermediate trailing edge flap 2Ialso hinged. abeutline` il and located intermediate the beforementionedflap I8 and aileron I9. The said flaps: Ihandizb are adapted to beactuated by means of conventional controls in the pilots compartmentthrough suitable coupling means such as cables or push rods so that theymay be raised or lowered simultaneously with corresponding flaps on theopposite wing semispan and aileron flaps 2I or each semispan wingtip-may be differentially operated in the conventional manner.

The present invention is applicable to wings having substantiallyany'desired plan form varyingffromrectangular to'sharply tapered andthose cf tapered plan form may vary in chord and thickness in accordancewith a curvilinear or straight line function of the span. In otherWords, the invention is applicable to rectangular wings or to taperedwings in which the leading and trailing edges may be either straight orcurved. The airfoil section may in either case be constant throughoutthawing span orA it'may vary from point topoint along the span, usuallyin such a manner as to undergo a uniform transition from one givenairfoil-section at the root to another airfoil section at the tip andthis transition of airfoil section may be employed to advantage in thepresent invention` tov achieve aerodynamic or geometrical twistV orwash-out as hereinafter more fully described.

For the` purpose of simplified illustration and description of theVfundamentalpprinciples and means of attaining the objects' of thepresent invention reference is now made mainly to Figures 7 to 1l bymeans of4 which the manner Aof accomplishing the prf-:sent invention isillustrated,` and theV following approximate analysis of the aerodynamiccharacteristics" of'this wing wiil-serve to illustrate the fundamentalprinciples upon whichthe present invent-ion is; based. Reference isfirst made toY Figures 7A andl '7B in which a plan view andA approximateloading diagram respectively is shown of the semispan of a typical wingof conventional tapered plan form without sweep-back or twist;

Assuming the airfoil of Y Figure 7A to be in flight under conditionsl ofa given constant air speed the available lift exerted on anyrsmallincrement or spanwise sectional` element As of the wing will beA avalue'whch is directly proportional to the product of the chorddimension I by the coefficient OfltCLL of that element, and the totallift for thel seinispan` will be equal to the summation of the lift onall such increments or spanwise elementsthroughout the length of thewing semispan S.-

ifor example, theA airfoil section profile and the angle of incidenceisxthesame throughout the length of the span S,`then the-coefficient oflift, Ci. of the wing. neglectingvend effects will be the same value orconstant throughout the length of 'the span, or 1C'L=K and in a wing ofthe tapered plan form shown-and in which the chord length is as beforedescribed, a straight line function of the span.

Where l is the chordlength of any given chord elements, As at a distancefrom the tip, la is the chord length at the tip section and a, is therate of taper. Since the'lift `on any'given spanwise element of thevwing is proportional to the coecien't of lift multiplied by the chordof that element the left distribution neglecting end effects and otherminor factors will then be substantially uniform throughout the span andthis type of loading may be graphically represented in the mannerofFigure 7B.

The span loading is,l therefore, seen to be a straight line functionwhich has a slope which is proportional to the said `variation a, inchord taper from Wing tip toroot. Now, if the angle of attack ofA thisairfoil is varied throughout its practical range and within thesubstantially straight portion of the lift coefcient-angle of attackcurve the corresponding CL Variations will resultin shifting the .spanloading line up or down as the angle of `attack is increased ordecreased as llustratediby lines 3c, 3l` and 32 in Figure 7B. l

The `resulting areas under each of these loading lines will beproportional to the `total lift of the semispan of the wing and thecenters of gravity of all of the areasthus included under these linesare found to fall upon the common vertical line all which isperpendicular to the span. From this analysis it is apparent that anairfoil where the angle of incidence is constant throughout the span andwhether or not it has taper, the effective spanwise location of thecenter of pressure remains substantially at a constant distance from theroot section and does not move spanwise as the angle of attack isvaried. ISince with the usual cambered airfoil section the chordwisevcenter of pressure movement is forward with increased angles 'of `attackand rearward with decreased angles of attack, under the beforementionedconditions the wing is manifestly unstable.

Now, if the wing is given either geometrical or aerodynamic twist orwashout which varies, for example, as an inverse straight line functionfrom root to tip section, the Ci. of the wing elements for any givenangle Vof incidence of the wing will correspondingly vary substantiallyuniformly from root to tip in accordance with the equation i where CLTis the coefcient of lift at the tip section, b is a constant numericallyequal to the uniform rate of washout along the semispan and is thedistance of the spanwise element from the wing tip, but the tracerof thespan loading line will then be a function of the combined factors ofwing taper and washout, or

and since this is a second power exponential equation the span loadinglines will under these conditions be curved and therefore as the overallangle of attack of the wing carrying washout is varied as by rotation onits spanwise axis, the span loading lines will no longer move up or downin the manner illustrated in Figure 7B to include areas, the loci'of thecenters of gravity of which Vfall on a vertical straight line shown at34 but the slopes of the lines will be modified by the non-uniformity ofthe lift coefficient 'throughout the span to result in loading lineswhich appear curved as illustrated at 35, 36, 3l and 38 in which casethe spanwise locations or centroids of the centers of gravity of theareas included under the curved loading lines no longer fall at a fixedspanwise location but fall at variable spanwise locations as indicatedby the lines 39, 40, 4I and 42 in Figure 10. Now, by projecting thesecentroid lines to their points of intersection with their correspondingcenter of pressure lines as shown in Figure 8A as obtained from the dataof Figure 9 the coordinates of the center of pressure for the variousangles of attack may be established as shown `in Figure 8A and the locusof such center of pressure travel is seen to form a curved line as shownat 45 havving a negative slope with respect to the spanwise axis of theairfoil as more fully` explained `hereinafter. i

Referring now primarily to Figures 8 to 11 and for purposes ofillustrating the method of accomplishin'g the present invention, thefollowing example is given. The wing is assumed to have a taper ratio of3 to 1, a uniform twist or washout of 6 from the root plane of symmetrycenter` line Il to the tip (or an angle of attack differential of onedegree per station) and an airfoil section profile uniform throughoutthe span and to havethe lift coefiicient and center of press-urecharacteristics as given in Figure 9.

The said loading curves may be constructed by dividing the semi'span ofthe wing into `any suitable number of stations such as for example bystation lines Nos. 0 to 6, as shown in Figure 8A and computing theloading for eachstation line section of the airfoil as a value equal tothe product of the C'L for that station times the chord length andplotting the resultant values as corresponding ordinates foreach of suchstations. The values for CL for each station may be obtained from anysuitable airfoil characteristic curve and in the present case, as beforestated, these values were taken by way of example from the data ofFigure 9 which show the relationship between angle of `attack and`coefficients of lift and center of pressure for a typical airfoilsection approximating in profile that of the ClarkY.

First taking the values of CL from the 0 curve which indicates thecharacteristic of the airfoil with the flap in neutral or undeflectedposition and with the root section at station No. O as` sumed to be setat an angle of attack of +1, the

products of coefficient of lift and chord length for each are plotted ateach station taking into account the twist in the angle of incidencethroughout the span and loading curve 35 constructed through the pointsthus established. The same operations are repeated for any desirednumber of angles of attack; for example curves 35, 36, 31 and38 are forarbitrarily chosen root angles of attack of 1, 6, 11` and 16respectively. Next the respective spanwise location of the center ofgravities of the areas included under the curves 35-38 are projectedfrom centroid lines 39 to42 onto the wing plan form` of Figure 8A to thepoint of intersection with the corresponding chordwise centers ofpressure location lines,` and curve 45 constructed through such point toestablish the locus of the center of pressure travel as beforementioned.The beforemenenea-oas tioned'f. intersecting center.v of pressureflines-.imay

be:A approximately. located by plotting@ on. Fig-ure.: lA-k at.. eachstation` line the;.corresponding.;locaition' ofgthecenter ofipressurefor thesectioinzpro:l file andl angleJ of. attack; at; such'. stationspointsand; constructing spanwise curves'.throughA theses points. Thesecurves. areL found-.i to be substanetiallycstraightzlines. exceptV for`the extremely; low. angles-.ofi attack` condition'.

Nextrthe, semispan Yloading.:curveslill-:to 532 may: be similarlyconstructed for the condition where the u:Hap which `extends over theinnermoststhree station intervalsrof the wing; isA lowered toisomefgiven'angle-such as forexamplelf? and thecor-re'-Y sponding lift values:under such condition. corre. spondingto various angles ofattaclnatthefseveralstations,` are obtained from .the values-.of CLtaken from the 45t/curve ofL-lligureQ,A In this-condition, since thereis-V a marked: discontinuity inA the chordw-isef center of. pressurelocati-oni, fof the outertip section which carriesfnotdepressedflan andthe root section which carries the depressed flap, thecorrespondingcenter: oi--pressurelines for.` thesev portions. of. thewing areseparately. established-from data of Figure` 9; andthe-spamrwiseflocations. ofthe centers. of gravity of the core4 responding.portions of-.the loading.4 curves-z as respeotivelyi indicated byllines55 toi'andlltogltare projected to intersectiontherewith to estab.-lishnseparate loci-of: center.. of pressureztravel i or each iquarter;span .portion .as-.shownaat 63 and z (it: inililirgureA. Now; .byestablishing..thweighted resultant center of; pressure: points` between:the locus` curves G3 and 64 the resultant; linea-6.5;' is establishedwhich; represents. the. locus.v of; the dernier:` of pressure. travel 4for the condition Where the flap l Bg is ;1owered :45,;

The-locus line Glilis4 similarly-` established for. the conditionwherethe-ap israised 5.-?'by-. employing the center,- of gravity lines;v80; to'- 8-3..

The wingof `Eigure-A without sweeprloack-` and l'iavinggtheIthus-:established center ot'pressure loci lg-liandsl is ohviousls7unstable since it; is rape parenti. that the; chordwise; center of.pressuretravel .isalways in .a direction-tending to .increase rather,than decrease-any-` angle of attack'. displacement-which may occur:

Now, iny order to attain longitudinal, stability, the-longitudinalcenter ofgravity line-must-,pass through the; centery of pressurelocus-.lineA aty an angle as shown at 61 and through the point- 69. oncurve-45 correspondingto the: angle. of. inci dence atwhich it-isdesirable to have-theairplane in trim so that as the angle ofincidence-1 is increased. above the; angleof trim the centerv ofpressure passes rearward .of the line B'L-toproduce a` counteractingcouple about. line- 6'!- tendingto increase the diving momentandreturnthe .airplane. to ,the trim` angle. Likewise, whenthe. 'angleofincidenceis reduced vto .a value belowthe .trim angle, the centerofpressure mustmove forward oncurve ,45 with .respect to. the. center. of.gravity line. E] to produce a oounteracting couple tending toincreasethe stalling momentand return` the airplane to the trim angle. The angle.between line 6T andthe lateral axis1 line 'llrindicatesthe approximateanglev ofA sweep-back necessary to be impartedto the airfoil ofFigure-SAto attain thisA condition. Therefore-,- if the elementarychordsectionsofthe wingfoi"JA Figure 8A are each moved uniformlybackward intheir own. planes without reducing the-span until;a..sweep;.back ofapproximatelyiqiidegrees is attained,.the longifltudinalf center of gravity-fposition line (il` wil-libe brought about toa` positionas shownzbyfthe rectilinean line at Gityirn. Figure:-` SB;extending; laterally; ati; the: same,l angl-ei through; the@ locus;line;v .'Il'ie2 locus-.linea williin: be.- in effect rotated to a newposition as shnwm` at'fll-,suchthat as the angle of` attack;l of thewing is decreased the center of pressure moves inward and forward: andasjthe angleof: attach-:is .increased` tha-centen 0inressurenmoresffoutward; andxzreareward, thus; attaining; ai condition.;of. automatic longitudinalstabilitx..

The-,centercof pressuralocus line; IZfintersects the; center:` o fgravityg line:- 68=z at;v point: 1.4- which correspondstoL aewing. angle.of attack-of; with the.: flap deflected .45?" downward,..` The`airplane,- thereforeimay-be; trimmeduat this; high-.angle of. attackylowering-i: the, flap.V to.. 45 since: under this condition thenitchingg moments.Y arefzero... Intermediate'flapangles will eiectztrinrat ang-les intermediateitheemaximum.angle .of 116;"y and the;trirnangleI of approximately 2?'as:;determined.by theipoi-ntiotintersection 12cofithe '.centenofzerayity line 681; through .the:1ocus;curve; 'H Ifj the` ap [8i is raisedito: some. anglegsuch: as.approximately 5? above-neutralithescenter o f: pressure will fall inlocus L line 13ibehindr; the-center. of f gravity:- line (i8 toieifectejnincreasain-.the diving moment. The angletoQwhichzthe: flap mustberaised-l-to obtain ,trim at,. say, .19 isgapproximately. indicated bythe ratio of the distance forward andaftiofthe center. of gravity. line:BR-Lbetweeni the curves.,l 13 and .1 I, andfthe approximatetrim anglesof'attackfor various flap settingsrcanz-be estimatedin this manner.;'Ihe'correspondingftrim anglesofattack forwarious-iiap .settings maybeplotted as rshown atline 'I5120i*Fig-ure:9;

From theioregoing it'islapparent that an angle of sweep-backand lpartial'span flap -combination can. loe` attained'i which. willl allow: tl'iepitching moment/tome; controllediby: the-flap so that-.a stalling:vmom-entf can bef-effected by lowering the 'ap'andfa diving momenteffected b'y-raising the` i'ap.

Fig-ure 1-1 illustrates diagrammatically.` and qualitatively what*effeeti variationof the. spanwise lengthof the -flap---hasn on the:pitching: moment` it is capable-of-e-fecting. Accordingly,- asthe-ap-is-increased in-spanwise=-length, a given downward deflection ofthe-flap-producesy progressively greater-A stallingmomentsL untilanoptimum-- length isv reached-i somewhere nearthe spanwise positiorr o-fthe mean-\ aerodynamic chord-as inditatedd at lifonv the curve.- Furtherlengthening ofi the flap resultsl indiminishing pitching moment' forthe-f samedeiieetion until a length of* approximately 75% of the-spanisreached as indicated at TI on the curve where-the flap is ineffectivet'opaltentheilongitudinal pitching moment. Atthis .point the flaps ofsuch spanwise length may,beglowe1'ed-jto-increasethe maximum eiectivecoefficient. of.` lift" offtlie airfoi-l withoutprodncmg either divingor stal1ingjmo ment. in. the airplane.. Such an arrangementwouldfbedesirable in-.somelcasesfwhere flaps are desired topeemployedasmigh lift..de.vicesvor brakes without imposing the large addeddi-Ving moments usuallyl associatedlwitl'i i their. use. For example,itis desirablein some casesin either the conventional :airplaneonin.these-.called ,forward tail. .airplane ,to-avoid ,the high..balanclngdoading imposed-.- upon f thefnorizontal-fstabilizer. when-theapsare loweredfor-aextended Ewith-the-:attendant rearward, movement otthecenter ot; pressn-re by employingsufiieient-,sweepslcack.andiroet-.sec tion apssot such-lengtnas'tominimize c-,or-:elim-ieV nate the pitching moment eiected thereby, inthe manner of this invention.

With further reference to Figure 11, further extension of the spanwiselength of the flap past point 'l1 up to full span flaps, results inpitching moments which are opposite in sense to that produced by thebeforementioned shorter ilaps on the sweep-back wing. Y i

If desired, three separate iiaps may be employed as illustrated at I8,I9 and 2| in Figure 6 and the combined length of iiaps I8 and 2| may bemade such as to extend to the length corresponding to point I asindicated in Figure 11, whereby their simultaneous use can be employedto eiiect an overall increased maximum coefficient of lift withoutappreciably changing the trim angle of the airplane. Flap I9 may beemployed separately for imposing the desired changes in pitching momentfor maneuvers, and naps I9 adjacent the wing tips may be operateddifferentially in the conventional manner of ailerons for affectinglateral control.

Either plain or split flaps as shown in Figures 4 and 5 may be employed.The split type of flaps are particularly suitable where the primaryobject is to obtain maximum coefcients of lift of the main wing with aminimum shift of the center of pressure. The plain flap as shown inFigure 4 is preferable ,particularly for the innermost flap such as napI8 where it serves as an elevator to provide longitudinal control forthe airplane. In such cases it is desirable to provide for raising theflap above the neutral position for which the split ap is not suited, toincrease the diving moment.

While the invention has been herein described as particularly applicableto the so-called tailless type of airplane, it is also applicable toeither of the conventional rearward tail or forward tail types. In theseother types of airplanes the l invention may be advantageously employedto reduce the stabilizing or balancing forces which from thelongitudinal plane of symmetry Without interference by the fuselage.This condition is possible in a low-wing monoplane where the split apsmay extend beneath the fuselage or in. a parasol type of high wingmonoplane where the wing and plain hinged aps are in the clear above thefuselage. However, in a mid-wing type of airplane as illustrated inFigures 1 to 3, the root portion of the wing semispan will be containedwithin the fuselage necessitating a slight shortening of the ilaps I 8.This will cause only a slight change in the control characteristicswithout departure from the principle of this invention.

The aps illustrated and employed in the computations of the presentinvention have a width of 0.20 of the wing chordbut flap widths varyingfrom 0.10 to 0.30 or greater have also been found suitable.

The foregoing is illustrative of a preferred method and embodiment ofthe invention and is not to be considered limiting since manymodications and adaptations may possibly be made by those skilled in theart within the scope of the claims.

I claim:

1. A tailless aircraft, comprising: a swept-back tapered wing includingwing semispans with washout of incidence and straight leading andtrailing edges from root to tip of each wing semispan, the ratio of theamount of taper to the degree of. washout being such that the locus ofapproximate points of resultant centers of pressure on each wingsemispan at operating angles of attack is an arcuate line extendinggenerally spanwise of the wing, and the degree of sweepback of the wingsemispans being of an amount to locate the center of pressure pointscorresponding to the angle of trim of each wing semispan on arectilinear line passing through the center of gravity of said aircraftso that the points on the arcuate line for increasing angles of attacklie rearwardly of said rectilinear line and outwardly toward the tip ofeach wing semispan and the points on the arcuate line for decreasingangles of attack lie forwardly of said rectilinear line and inwardlytoward the root of each wing semispan; trailing edge iiap means hingedto each wing semispan inwardly of the tip portion thereof; and means foreffecting angular displacement cf said flap means, the major portion ofsaid ap means being located to the rear of said rectilinear line,whereby downwardly displaced positions of said flap means cause forwarddisplacement of the locus of said resultant centers of pressure, therebyvarying simultaneously and inversely with respect to one another themaximum effective lift and center of pressure coefficients of said wingrelative to the center of gravity of said aircraft.

2. A tailless aircraft as dened in claim 1, in which the flap meansconsists of a flap extending spanwise of the wing adjacent the root ofeach wing semispan; and in which the aircraft further includes anaileron adjacent the tip portion of each wing semispan, and means foreffecting angular displacement of said ailerons.

3. A tailless aircraft as dened in claim 1, in which the iiap meanscomprises a pair of adjacent flaps on each wing semispan with one flapof each pair constituting an inner ap and extending spanwise from apoint adjacent the root of its associated wing semispan and with theother flap of each pair being disposed outwardly of its associated innernap.

4. A tailless aircraft as defined in claim 1, in which the nap meanscomprises a pair of adjacent iiaps on each wing semispan with one iiapof each pair constituting an inner flap and extending spanwise from apoint adjacent the root of its associated wing semispan and with theother ap of each `pair constituting 'an intermediate lap and beingdisposed outwardly of its associated inner flap; and in which theaircraft further includes an aileron disposed outwardly of saidintermediate flaps and adjacent the tip portion of each wing semispan,and means for effecting angular displacement of said ailerons.

WALTER. H. KORFF.

REFERENCES CITED The following references are of record in the le ofthis patent:

UNITED STATES PATENTS Number Name IDate D. 127,185 Northrop 1 May 13,1941 1,003,721 Dunne Sept. 19, 1911 1,600,671 Hill Sept. 21, 19261,780,813 Burnelli Nov. 4, 1930 1,987,050 Burnelli Jan. 8, 19352,130,958 Kramer Sept. 20, 1938 2,172,289 Munk Sept. 5, 1939 (Otherreferences on following page)

